Low noise turbine for geared gas turbine engine

ABSTRACT

A gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, a propulsor section including a propulsor, a turbine section including a first turbine and a second turbine, a compressor section driven by the turbine section, the compressor section including a first compressor and a second compressor, and a geared architecture driven by the first turbine. The propulsor is driven by the first turbine via the geared architecture. At least one stage of the turbine section includes an array of rotatable blades and an array of vanes. A ratio of the number of vanes to the number blades is greater than or equal to 1.55. A mechanical tip rotational Mach number of the blades is configured to be greater than or equal to 0.5 at an approach speed.

CROSS-REFERENCE TO RELATED APPLICATIONS

This application is a continuation of U.S. patent application Ser. No.16/833,782, filed Mar. 30, 2020, which is a continuation of U.S. patentapplication Ser. No. 13/970,670, filed Aug. 20, 2013, which claimspriority to U.S. Provisional Application No. 61/781,170, which was filedon 14 Mar. 2013 and is incorporated herein by reference.

BACKGROUND

This disclosure relates to the design of a lower noise gas turbineengine turbine.

Gas turbine engines are known, and typically include a fan deliveringair into a compressor. The air is compressed in the compressor anddelivered downstream into a combustor section where it is mixed withfuel and ignited. Products of this combustion pass downstream overturbine rotors, driving the turbine rotors to rotate.

Typically, there is a high pressure turbine rotor, and a low pressureturbine rotor. Each of the turbine rotors include a number of rows ofturbine blades that rotate with the rotor. Interspersed between the rowsof turbine blades are vanes.

The high pressure turbine rotor has typically driven a high pressurecompressor rotor, and the low pressure turbine rotor has typicallydriven a low pressure compressor rotor. Each of the compressor rotorsalso include a number of compressor blades that rotate with the rotors.There are also vanes interspersed between the rows of compressor blades.

The low pressure turbine or compressor can be a significant noisesource, as noise is produced by fluid dynamic interaction between theblade rows and the vane rows. These interactions produce tones at ablade passage frequency of each of the low pressure turbine rotors, thelow pressure compressor rotors, and their harmonics.

Historically, the low pressure turbine has driven both a low pressurecompressor section and a fan section. More recently, a gear reductionhas been provided such that the fan and low pressure compressor can bedriven at distinct speeds.

With the inclusion of a gear, low pressure turbine speeds haveincreased. Thus, to “cutoff” these turbines, vane-to-blade ratios mustbe higher than for turbines in a conventional engine.

SUMMARY

A gas turbine engine according to an exemplary aspect of the presentdisclosure includes, among other things, a turbine section including afan drive turbine, a compressor section driven by the turbine section, ageared architecture driven by the fan drive turbine, and a fan driven bythe fan drive turbine via the geared architecture. At least one stage ofthe turbine section includes an array of rotatable blades and an arrayof vanes. A ratio of number of vanes to the number blades is greaterthan or equal to about 1.55. A mechanical tip rotational Mach number ofthe blades is configured to be greater than or equal to about 0.5 at anapproach speed.

In a further non-limiting embodiment of the foregoing gas turbineengine, the vanes of the at least one stage are immediately upstream ordownstream from the blades.

In a further non-limiting embodiment of any of the foregoing gas turbineengines, the gas turbine engine is rated to produce 15,000 pounds ofthrust or more.

In a further non-limiting embodiment of any of the foregoing gas turbineengines, the at least one stage comprises a stage of a low pressureturbine.

In a further non-limiting embodiment of any of the foregoing gas turbineengines, the at least one stage comprises each stage of a low pressureturbine.

In a further non-limiting embodiment of any of the foregoing gas turbineengines, the gear reduction has a gear ratio of greater than about 2.3.

In a further non-limiting embodiment of any of the foregoing gas turbineengines, the fan delivers air into a bypass duct, and a portion of airinto the compressor section, with a bypass ratio defined as the volumeof air delivered into the bypass duct compared to the volume of airdelivered into the compressor section, and the bypass ratio beinggreater than about six (6).

In a further non-limiting embodiment of any of the foregoing gas turbineengines, the bypass ratio is greater than about ten (10).

In a further non-limiting embodiment of any of the foregoing gas turbineengines, the turbine section is a turbine section of a three-spooled gasturbine engine.

A turbine section of a geared gas turbine engine according to anexemplary aspect of the present disclosure includes, among other things,at least one stage having a ratio of vanes to blades that is greaterthan or equal to about 1.55. The blades are configured to operate at amechanical tip rotational Mach number that is greater than or equal toabout 0.5 at an approach speed.

In a further non-limiting embodiment of the foregoing turbine section,the vanes of the at least one stage are immediately upstream ordownstream from the blades.

In a further non-limiting embodiment of any of the foregoing turbinesections the geared gas turbine engine is rated to produce 15,000 poundsof thrust or more.

In a further non-limiting embodiment of any of the foregoing turbinesections, the at least one stage comprises a stage of a low pressureturbine.

In a further non-limiting embodiment of any of the foregoing turbinesections, the at least one stage comprises each stage of a low pressureturbine.

A method of expansion in a gas turbine according to another exemplaryaspect of the present disclosure includes, among other things, providingat least one stage of a turbine section of a geared gas turbine engine,the at least one stage having a ratio of vanes to blades that is greaterthan or equal to about 1.55. The mechanical tip rotational Mach numberis configured to be greater than or equal to 0.5 at the approach speed.

In a further non-limiting embodiment of the foregoing method, the atleast one stage comprises at least one stage of a low pressure turbine.

These and other features of this disclosure will be best understood fromthe following specification and drawings, the following of which is abrief description.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 shows an example gas turbine engine.

DETAILED DESCRIPTION

FIG. 1 schematically illustrates an example gas turbine engine 20 thatincludes a fan section 22, a compressor section 24, a combustor section26, and a turbine section 28. Alternative engines might include anaugmenter section (not shown) among other systems or features. The fansection 22 drives air along a bypass flow path B while the compressorsection 24 draws air in along a core flow path C where air is compressedand communicated to a combustor section 26. In the combustor section 26,air is mixed with fuel and ignited to generate a high pressure exhaustgas stream that expands through the turbine section 28 where energy isextracted and utilized to drive the fan section 22 and the compressorsection 24.

Although the disclosed non-limiting embodiment depicts a turbofan gasturbine engine, it should be understood that the concepts describedherein are not limited to use with turbofans as the teachings may beapplied to other types of turbine engines; for example a turbine engineincluding a three-spool architecture in which three spoolsconcentrically rotate about a common axis and where a low spool enablesa low pressure turbine to drive a fan via a gearbox, an intermediatespool that enables an intermediate pressure turbine to drive a firstcompressor of the compressor section, and a high spool that enables ahigh pressure turbine to drive a high pressure compressor of thecompressor section.

The example engine 20 generally includes a low speed spool 30 and a highspeed spool 32 mounted for rotation about an engine central longitudinalaxis A relative to an engine static structure 36 via several bearingsystems 38. It should be understood that various bearing systems 38 atvarious locations may alternatively or additionally be provided.

The low speed spool 30 generally includes an inner shaft 40 thatconnects a fan 42 and a low pressure (or first) compressor section 44 toa low pressure (or first) turbine section 46. The inner shaft 40 drivesthe fan 42 through a speed change device, such as a geared architecture48, to drive the fan 42 at a lower speed than the low speed spool 30.The high speed spool 32 includes an outer shaft 50 that interconnects ahigh pressure (or second) compressor section 52 and a high pressure (orsecond) turbine section 54. The inner shaft 40 and the outer shaft 50are concentric and rotate via the bearing systems 38 about the enginecentral longitudinal axis A.

A combustor 56 is arranged between the high pressure compressor 52 andthe high pressure turbine 54. In one example, the high pressure turbine54 includes at least two stages to provide a double stage high pressureturbine 54. In another example, the high pressure turbine 54 includesonly a single stage. As used herein, a “high pressure” compressor orturbine experiences a higher pressure than a corresponding “lowpressure” compressor or turbine.

The example low pressure turbine 46 has a pressure ratio that is greaterthan about five (5). The pressure ratio of the example low pressureturbine 46 is measured prior to an inlet of the low pressure turbine 46as related to the pressure measured at the outlet of the low pressureturbine 46 prior to an exhaust nozzle.

A mid-turbine frame 58 of the engine static structure 36 is arrangedgenerally between the high pressure turbine 54 and the low pressureturbine 46. The mid-turbine frame 58 further supports bearing systems 38in the turbine section 28 as well as setting airflow entering the lowpressure turbine 46.

The core airflow C is compressed by the low pressure compressor 44 thenby the high pressure compressor 52 mixed with fuel and ignited in thecombustor 56 to produce high speed exhaust gases that are then expandedthrough the high pressure turbine 54 and low pressure turbine 46. Themid-turbine frame 58 includes vanes 60, which are in the core airflowpath and function as an inlet guide vane for the low pressure turbine46. Utilizing the vane 60 of the mid-turbine frame 58 as the inlet guidevane for low pressure turbine 46 decreases the length of the lowpressure turbine 46 without increasing the axial length of themid-turbine frame 58. Reducing or eliminating the number of vanes in thelow pressure turbine 46 shortens the axial length of the turbine section28. Thus, the compactness of the gas turbine engine 20 is increased anda higher power density may be achieved.

The disclosed gas turbine engine 20 in one example is a high-bypassgeared aircraft engine. In a further example, the gas turbine engine 20includes a bypass ratio greater than about six (6), with an exampleembodiment being greater than about ten (10). The example gearedarchitecture 48 is an epicyclical gear train, such as a planetary gearsystem, star gear system or other known gear system, with a gearreduction ratio of greater than about 2.3.

In one disclosed embodiment, the gas turbine engine 20 includes a bypassratio greater than about ten (10:1) and the fan diameter issignificantly larger than an outer diameter of the low pressurecompressor 44. It should be understood, however, that the aboveparameters are only exemplary of one embodiment of a gas turbine engineincluding a geared architecture and that the present disclosure isapplicable to other gas turbine engines.

A significant amount of thrust is provided by the bypass flow B due tothe high bypass ratio. The fan section 22 of the engine 20 is designedfor a particular flight condition—typically cruise at about 0.8 Mach andabout 35,000 feet. The flight condition of 0.8 Mach and 35,000 ft., withthe engine at its best fuel consumption—also known as “bucket cruiseThrust Specific Fuel Consumption (“TSFC”)”—is the industry standardparameter of pound-mass (lbm) of fuel per hour being burned divided bypound-force (lbf) of thrust the engine produces at that minimum point.

“Low fan pressure ratio” is the pressure ratio across the fan bladealone, without a Fan Exit Guide Vane (“FEGV”) system. The low fanpressure ratio as disclosed herein according to one non-limitingembodiment is less than about 1.50. In another non-limiting embodimentthe low fan pressure ratio is less than about 1.45.

“Low corrected fan tip speed” is the actual fan tip speed in ft/secdivided by an industry standard temperature correction of [(Tram °R)/(518.7° R)]^(0.5). The “Low corrected fan tip speed,” as disclosedherein according to one non-limiting embodiment, is less than about 1150ft/second.

The example gas turbine engine includes the fan 42 that comprises in onenon-limiting embodiment less than about twenty-six (26) fan blades. Inanother non-limiting embodiment, the fan section 22 includes less thanabout twenty (20) fan blades. Moreover, in one disclosed embodiment thelow pressure turbine 46 includes no more than about six (6) turbinerotors schematically indicated at 34. In another non-limiting exampleembodiment the low pressure turbine 46 includes about three (3) turbinerotors. The example low pressure turbine 46 provides the driving powerto rotate the fan section 22 and therefore the relationship between thenumber of turbine rotors 34 in the low pressure turbine 46 and thenumber of blades in the fan section 22 disclose an example gas turbineengine 20 with increased power transfer efficiency.

The use of the gear reduction between the low speed spool 30 and the fan42 allows an increase of speed to the low pressure turbine 46. In thepast, the speed of the low pressure turbine 46 and the low pressurecompressor 44 has been somewhat limited in that the fan speed cannot beunduly large. The maximum fan speed is at its outer tip, and in largerengines, the fan diameter is much larger than it may be in smaller powerengines. However, the use of the gear reduction has freed the designerfrom limitation on the speeds of the low pressure turbine 46 and the lowpressure compressor 44 speeds caused by a desire to not have unduly highfan speeds.

In geared gas turbine engines, such as the engine 20, a careful designbetween the number of vanes and blades in the low pressure turbine 46,and the mechanical tip rotational Mach number of the low pressureturbine 46 can be selected to reduce turbine noise through the use ofthe mechanism referred to as “cutoff.” This “cutoff” mechanism occurswhen the vane-to-blade ratio is selected such that the fundamental bladepassage tone is prevented from propagating to the far-field. Thismechanism has been used previously in non-geared engines, which have lowpressure turbines that operate at low tip Mach numbers, typically nogreater than 0.5. However, “cutoff” has not been used in geared engines,such as those described herein, which have low pressure turbines thatoperate at high tip Mach numbers, typically greater than 0.5. On gearedengines with such turbines, the “cutoff” mechanism requires a largervane-to-blade ratio than it would on non-geared engines.

The mechanical tip rotational Mach number, M_(tip), is generally definedas:

$M_{tip} = {\frac{\pi}{720c}DN}$wherein N is a rotor rotational speed in revolutions per minute, c isthe local speed of sound in feet per second and D is the local tipdiameter in inches.

The mechanical tip rotational Mach number for any blade row may becalculated in this manner.

Although described with reference to the two-spool engine 20, therelationship between the number of vanes and blades in the low pressureturbine 46, and the mechanical tip rotational Mach number of the lowpressure turbine 46 may be applicable to three-spool direct driveengines or three-spool engines having a gear reduction as well.

In the example engine 20, a ratio of the number of vanes to blades in astage of the low pressure turbine is greater than or equal to R_(A). Inthis example, a mechanical tip rotational Mach number of the blade ofthe low pressure turbine is greater than or equal to M_(A) at approachspeed. In the example engine 20, R_(A) is about 1.55 and M_(A) is about0.5. This novel design will result in reduced low pressure turbine noisebecause at least one stage of the low pressure turbine is “cutoff” atits rotor blade passing frequency.

The stage including the vanes and blades greater than or equal to R_(A),can be any stage of the low pressure turbine 46.

The stage may also be a stage of the high pressure turbine 54, or, ifpresent, an intermediate pressure turbine. In a high or intermediatepressure turbine example, R_(A) may be greater than or equal to 1.55.

It is envisioned that all of the stages in the low pressure turbine 46(or high pressure turbine 54 or, if present, an intermediate pressureturbine) would include a ratio of vanes to blades that is greater thanor equal to R_(A). However, this disclosure may also extend to turbineswherein only one of the stages has a ratio of vanes to blades that isgreater than or equal to R_(A). This disclosure also extends to turbineswherein more than one, but less than all, of the stages has a ratio ofvanes to blades that is greater than or equal to R_(A).

The mechanical tip rotational Mach number is measured at engineoperating conditions corresponding to one or more of the noisecertification points defined in Part 36 of the Federal AirworthinessRegulations. More particularly, the rotational speed may be taken as anapproach certification point as defined in Part 36 of the FederalAirworthiness Regulations. For purposes of this application and itsclaims, the term “approach speed” equates to this certification point.

The disclosed examples are most applicable to jet engines rated toproduce 15,000 pounds (66,723 N) of thrust or more.

Although an embodiment of this invention has been disclosed, a worker ofordinary skill in this art would recognize that certain modificationswould come within the scope of this invention. For that reason, thefollowing claims should be studied to determine the true scope andcontent of this invention.

What is claimed is:
 1. A gas turbine engine comprising: a propulsorsection including a propulsor having a plurality of propulsor blades; aturbine section including a first turbine and a second turbine; acompressor section driven by the turbine section, the compressor sectionincluding a first compressor and a second compressor; a gearedarchitecture driven by the first turbine; a first shaft that connects aninput of the geared architecture to the first turbine; a second shaftthat connects the second compressor and the second turbine, and whereinthe first and second shafts are concentric and are rotatable via bearingsystems about an engine axis; wherein the propulsor is driven by thefirst turbine via the geared architecture, and the second turbine drivesthe second compressor; wherein at least one stage of the turbine sectionincludes an array of rotatable blades and an array of vanes, and thearray of vanes of the at least one stage are immediately upstream ordownstream from the array of blades; wherein a ratio of the number ofvanes to the number of blades of the at least one stage is greater thanor equal to 1.55; wherein a mechanical tip rotational Mach number of thearray of blades is greater than or equal to 0.5 at an approach speed,the approach speed taken at an approach certification point as definedin Part 36 of the Federal Airworthiness Regulation; and wherein the gasturbine engine is rated to produce 15,000 pounds of thrust or more. 2.The gas turbine engine as recited in claim 1, wherein the gearedarchitecture is an epicyclic gear train.
 3. The gas turbine engine asrecited in claim 2, wherein the first turbine includes a plurality ofstages, and the at least one stage comprises at least one stage of theplurality of stages of the first turbine.
 4. The gas turbine engine asrecited in claim 3, wherein the first compressor includes a plurality ofstages.
 5. The gas turbine engine as recited in claim 4, wherein thegeared architecture has a gear ratio of greater than 2.3.
 6. The gasturbine engine as recited in claim 5, wherein the first turbine includesthree (3) turbine rotors.
 7. The gas turbine engine as recited in claim6, wherein the second turbine includes two stages.
 8. The gas turbineengine as recited in claim 7, further comprising a pressure ratio ofless than 1.50 across the propulsor blade alone.
 9. The gas turbineengine as recited in claim 8, wherein the first turbine drives the firstcompressor and an input of the geared architecture.
 10. The gas turbineengine as recited in claim 9, wherein the first compressor has a greaternumber of stages than the second turbine.
 11. The gas turbine engine asrecited in claim 10, wherein the propulsor has less than twenty-six (26)propulsor blades.
 12. The gas turbine engine as recited in claim 11,wherein: the propulsor has a low corrected tip speed of less than 1150ft/second; and the first turbine includes an inlet, an outlet and apressure ratio of greater than 5, the pressure ratio being pressuremeasured prior to the inlet as related to pressure at the outlet priorto an exhaust nozzle.
 13. The gas turbine engine as recited in claim 12,wherein the geared architecture is a star gear system.
 14. The gasturbine engine as recited in claim 13, wherein the first turbineincludes no more than six (6) turbine rotors.
 15. The gas turbine engineas recited in claim 14, wherein the at least one stage comprises onlyone stage of the first turbine.
 16. The gas turbine engine as recited inclaim 14, wherein the at least one stage comprises more than one stageof the plurality of stages of the first turbine.
 17. The gas turbineengine as recited in claim 16, wherein the pressure ratio is less than1.45 across the propulsor blade alone.
 18. The gas turbine engine asrecited in claim 17, wherein the propulsor has less than twenty (20)propulsor blades.
 19. The gas turbine engine as recited in claim 18,wherein the at least one stage comprises less than all of the stages ofthe first turbine.
 20. The gas turbine engine as recited in claim 18,wherein the at least one stage comprises all of the stages of the firstturbine.
 21. The gas turbine engine as recited in claim 20, wherein theturbine section includes a mid-turbine frame between the first turbineand the second turbine, the mid-turbine frame supports bearing systemsin the turbine section, and the mid-turbine frame includes vanes in acore flow path.
 22. The gas turbine engine as recited in claim 12,wherein the geared architecture is a planetary gear system.
 23. The gasturbine engine as recited in claim 22, wherein the first turbineincludes no more than six (6) turbine rotors.
 24. The gas turbine engineas recited in claim 23, wherein the at least one stage comprises onlyone stage of the first turbine.
 25. The gas turbine engine as recited inclaim 23, wherein the at least one stage comprises more than one stageof the plurality of stages of the first turbine.
 26. The gas turbineengine as recited in claim 25, wherein the pressure ratio is less than1.45 across the propulsor blade alone.
 27. The gas turbine engine asrecited in claim 26, wherein the propulsor has less than twenty (20)propulsor blades.
 28. The gas turbine engine as recited in claim 27,wherein the at least one stage comprises less than all of the stages ofthe first turbine.
 29. The gas turbine engine as recited in claim 27,wherein the at least one stage comprises all of the stages of the firstturbine.
 30. The gas turbine engine as recited in claim 29, wherein theturbine section includes a mid-turbine frame between the first turbineand the second turbine, the mid-turbine frame supports bearing systemsin the turbine section, and the mid-turbine frame includes vanes in acore flow path.